US8662828B2 - High pressure turbine of a turbomachine with improved assembly of the mobile blade radial clearance control box - Google Patents

High pressure turbine of a turbomachine with improved assembly of the mobile blade radial clearance control box Download PDF

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Publication number
US8662828B2
US8662828B2 US12/994,969 US99496909A US8662828B2 US 8662828 B2 US8662828 B2 US 8662828B2 US 99496909 A US99496909 A US 99496909A US 8662828 B2 US8662828 B2 US 8662828B2
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Prior art keywords
shock absorbing
high pressure
pressure turbine
annular
control box
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US12/994,969
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US20110076135A1 (en
Inventor
Alain Dominique Gendraud
Delphine Roussin-Leroux
Jean-Luc Le Strat
Pascal Tatiossian
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Safran Aircraft Engines SAS
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SNECMA SAS
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Assigned to SNECMA reassignment SNECMA ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: GENDRAUD, ALAIN DOMINIQUE, LE STRAT, JEAN-LUC, ROUSSIN-LEROUX, DELPHINE, TATIOSSIAN, PASCAL
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Assigned to SAFRAN AIRCRAFT ENGINES reassignment SAFRAN AIRCRAFT ENGINES CHANGE OF NAME (SEE DOCUMENT FOR DETAILS). Assignors: SNECMA
Assigned to SAFRAN AIRCRAFT ENGINES reassignment SAFRAN AIRCRAFT ENGINES CORRECTIVE ASSIGNMENT TO CORRECT THE COVER SHEET TO REMOVE APPLICATION NOS. 10250419, 10786507, 10786409, 12416418, 12531115, 12996294, 12094637 12416422 PREVIOUSLY RECORDED ON REEL 046479 FRAME 0807. ASSIGNOR(S) HEREBY CONFIRMS THE CHANGE OF NAME. Assignors: SNECMA
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/14—Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing
    • F01D11/20—Actively adjusting tip-clearance
    • F01D11/24—Actively adjusting tip-clearance by selectively cooling-heating stator or rotor components
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/30—Retaining components in desired mutual position
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/96—Preventing, counteracting or reducing vibration or noise

Definitions

  • This invention relates to a high pressure turbine in a turbomachine such as an aircraft turbojet or turboprop.
  • a high pressure turbine in a turbomachine comprises at least one stage containing a distributor formed from an annular row of fixed guide vanes and a bladed wheel mounted free to rotate on the downstream side of the distributor in a cylindrical or tapered and truncated cone shaped assembly of ring sectors placed circumferentially end to end. These ring sectors comprise means of attachment to an annular support at their forward and aft ends, and the annular support is fixed to a turbine external casing.
  • FIGS. 1 and 1A show one example assembly of the control box on the external casing of the turbomachine high pressure turbine. These figures show a wall 400 of the control box 40 fixed to the external casing 22 of the turbine 10 at two diametrically opposite points through a threaded ring 4 .
  • the inventors observed that during operation of the turbomachine equipped with the high pressure turbine, vibrations occur in the control box that can cause damage to its attachment points. There is a risk of cracks occurring at its attachment points.
  • the purpose of the invention is to disclose a solution to prevent disturbing loads at the attachment points of the control box to the external casing during operation of the high pressure turbine of a turbomachine.
  • the purpose of the invention is a high pressure turbine for a turbomachine comprising:
  • control box that can be used in the invention is disclosed in the embodiment described in patent application FR 2 865 237. Therefore the entire content of this prior application is included within this application.
  • the energy of the vibrations of the box generated by excitation modes of the turbomachine is dissipated by a combination of friction at the axial bearing and braking of the control box due to bending of the additional annular element.
  • the shock absorbing element according to the invention disturbs the development of harmful vibration modes.
  • the annular shock absorbing element is a metal section made by machining or forming a metal plate.
  • the geometric shape of the annular shock absorbing element is composed of a continuous crown fixed to the annular support and extended by a plurality of identical, equally spaced blades at an inclination from the crown, the curved end of which forms the pressure bearing with the upstream side of the box.
  • the number of blades in the annular shock absorbing element will be equal to a multiple of eighteen. Studies have shown that this choice, for example using seventy-two blades uniformly distributed around the circumference for a 0.680 m diameter have been perfectly satisfactory.
  • control box and the annular shock absorbing element are made from the same material.
  • anti-wear material is inserted in the bearing area between the shock absorbing element and the upstream side of the box so reduce wear of the shock absorber or the box due to friction.
  • a layer of anti-wear material is preferably deposited on the upstream side of the box in the bearing area with the shock absorbing element.
  • the annular shock absorbing element is composed of at least two angular sectors fixed end to end forming the complete annular shape of the shock absorber.
  • the shock absorbing element is preferably composed of two, six or eighteen angular sectors fixed end to end forming the complete annular shape of the shock absorber.
  • the annular shock absorbing element is fixed to the annular support through screws that also attach the axial spacer stops. These parts are usually called stop plates.
  • the metal section may comprise at least one angular sector of a continuous crown prolonged by a plurality of identical blades, uniformly spaced and inclined relative to the crown sector and the end of which is curved.
  • the invention also relates to a turbomachine comprising a high pressure turbine like that described above.
  • FIG. 1 is a diagrammatic cross-sectional view of a turbojet high pressure turbine at the attachment points of the control box to the external casing,
  • FIG. 1A is a detailed view of FIG. 1 showing an attachment zone of the control box attached to the external casing
  • FIG. 2 is a partial diagrammatic half longitudinal sectional view of a turbojet high pressure turbine according to the invention
  • FIG. 3 is a detailed perspective view of a shock absorbing element according to the invention.
  • FIG. 4 is a partial perspective sectional view of a high pressure turbine according to the invention made at the attachment points of the control box to the external casing.
  • FIG. 1 diagrammatically shows part of a turbomachine such as an aircraft turbojet or turboprop comprising a high pressure turbine 10 arranged on the downstream side of a combustion chamber 12 and the upstream side of a low pressure turbine 14 of the turbomachine.
  • a turbomachine such as an aircraft turbojet or turboprop comprising a high pressure turbine 10 arranged on the downstream side of a combustion chamber 12 and the upstream side of a low pressure turbine 14 of the turbomachine.
  • the combustion chamber 12 comprises an external wall of revolution 50 connected at its downstream end to a radially internal end of a tapered and truncated wall 58 that comprises a radially external annular flange 60 at its radially external end used to attach it to a corresponding annular flange 62 on an external casing 64 of the chamber.
  • the high pressure turbine 10 comprises a single turbine stage in which there is a distributor 16 formed of an annular row of fixed guide vanes, and a bladed wheel 18 fitted free to rotate on the downstream side of the distributor 16 .
  • the low pressure turbine 14 comprises several turbine stages, each of these stages also comprising a distributor and a bladed wheel, only the distributor 47 of the upstream low pressure stage being visible in FIG. 1 .
  • the bladed wheel 18 of the high pressure turbine rotates inside an approximately cylindrical assembly of ring sectors 20 placed circumferentially end to end and suspended from an external casing 22 of the turbine through an annular support 24 .
  • This annular support 24 comprises means 26 of attaching the ring sectors 20 at its internal periphery, and comprises a wall 28 that extends outwards and in the upstream direction and that is connected to a radially external flange 30 for attachment of the external casing 22 of the turbine at its radially external end.
  • the flange 60 is axially inserted between the flange 30 and the flange 62 of the external casing 22 of the turbine and is axially clamped between these flanges by appropriate screw-nut type means 7 .
  • the annular support 24 comprises two upstream and downstream radial annular walls 34 , 36 respectively at its internal periphery, that are connected to each other through a cylindrical wall 38 .
  • the radial walls 34 , 36 comprise cylindrical rims 340 , 360 facing downstream at their radially internal ends, and these rims 340 , 360 cooperate with circumferential hooks 201 , 202 fitted at the downstream and upstream ends of the ring sectors 20 .
  • An annular locking device 46 with a C-shaped section is axially engaged at the downstream end on the downstream cylindrical rim 360 of the support and on the downstream hooks 201 of the ring sectors 20 to lock the assembly.
  • the wall 28 of the annular support 24 and the tapered and truncated wall 58 of the chamber define an annular containment 80 supplied with ventilation and cooling air through orifices 82 formed in the tapered and truncated wall 58 .
  • Orifices not shown are formed in the upstream radial wall 34 of the annular support 24 to create fluid communication between the containment 80 and an annular cavity 86 to cool ring sectors 20 delimited on the outside by the cylindrical wall 28 of the annular support.
  • the external wall 66 of the distributor comprises an annular groove 74 at each of its upstream and downstream ends, opening up in the radial direction to the exterior.
  • Annular sealing packing 76 is housed in these grooves 74 and cooperates with cylindrical ribs 78 formed on the tapered and truncated wall 58 and on the upstream radial wall 34 of the annular support 24 respectively, to prevent gas from passing radially outwards from the turbine flowstream through the external wall 66 , and conversely to prevent air from passing radially inwards from the containment 80 into the turbine flowstream.
  • the radial clearance between the tips of the mobile blades of the bladed wheel 18 and the ring 20 also has to be minimised to increase the efficiency of the turbine.
  • This device D comprises a circular control box 40 surrounding the fixed ring 20 , and more precisely the annular support 24 .
  • control box 40 is designed either to cool or heat the upstream rib 240 and the downstream rib 242 of the annular support 24 by blowing air on these blades.
  • the annular support 24 contracts or expands as this air comes into contact with it, which reduces or increases the diameter of the fixed ring segments 20 of the turbine so as to vary the clearance at the blade tips of the bladed wheel 18 .
  • the control box 40 supports at least three annular air circulation headers 41 , 42 and 43 that surround the annular support 24 of the fixed ring assembly. These headers are at an axial spacing from each other and are approximately parallel to each other. They are arranged on each side of the side faces of each of the ribs 240 , 242 and they approximately match the shape of these ribs.
  • the control box 40 also comprises an air collection tube not shown to supply air to the air circulation headers 41 , 42 and 43 .
  • This air collection tube surrounds the headers 41 , 42 and 43 and supplies air to them through air ducts 44 .
  • such a control box 40 is composed of two half-shells clamped together and is fixed to the external casing 22 through threaded rings 45 at two diametrically opposite points ( FIG. 1 ).
  • the inventors have observed that during operation of the turbomachine comprising the high pressure turbine 10 as illustrated above, there can be a risk of cracks appearing at the attachment points 45 . They have demonstrated that this is due to the fact that the control box 40 is subjected to harmful vibrations that can cause damage at its attachment points 45 .
  • FIGS. 1 and 1A diagrammatically show elliptical contours representing the precise zones Z in which there is a risk of cracks appearing in the vicinity of the attachment holes 46 .
  • the invention mitigates this risk of cracks by implanting an annular element 5 with a predetermined flexibility inside the cavity delimited by the annular support 24 and the external casing 22 upstream from the control box 40 ( FIGS. 2 and 4 ).
  • This annular element 5 with a predetermined flexibility bearing at a given pressure thus forms a shock absorber for at least some of the vibrations of the control box 40 generated during operation of the turbine.
  • the damping thus provided according to the invention is a means of dissipating energy from vibrations of the box 40 generated during operation of the turbomachine by a combination of friction at the axial bearing at the other end 52 and braking of the control box 40 due to bending of the annular element between its ends 51 , 52 during operation of the turbomachine.
  • the shock absorbing element 5 improves energy dissipation and dynamic damping of the headers 41 , 42 and 43 controlling the radial clearance of the rotating blades of the bladed wheel 18 .
  • the shock absorbing element 5 thus provided avoids mechanical vibration loads from the control box 40 without needing to change the way in which it is attached to the external casing ( FIG. 4 ).
  • each angular sector forming the shock absorbing element 5 is a metal section obtained by forming a plate.
  • the geometric shape of the shock absorbing element 5 is composed of a continuous crown at the one end 51 fixed to the annular support and prolonged by a plurality of identical uniformly spaced blades 510 inclined relative to the crown at the one end 51 , the curved other end 52 of which forms the pressure bearing with the upstream part 401 of the box.
  • These blades 510 that create the pressure bearing on the upstream part of the box may for example be made from a continuous metal section by “saw cut” type machining diagrammatically represented by the space 53 between two consecutive blades 510 .
  • the number of blades 510 around the entire circumference can be varied by modifying the width of the saw cut made.
  • the number of blades in the shock absorbing element 5 is equal to a multiple of eighteen. For example, a number of blades equal to seventy-two would be desirable. Thirty-six or a hundred and forty-four blades could also be used.
  • the control box 40 and the shock absorbing element 5 are preferably made from the same material that could be a Hastelloy® X type alloy.
  • an anti-wear material in the bearing zone at the other end 52 between the shock absorbing element 5 and the upstream part 401 of the box 40 , to prevent premature wear of the control box 40 or the shock absorbing element 5 in mutual friction, and to improve energy dissipation by friction.
  • a Tribaloy® 800 or Tribaloy® 800 type alloy with CoCrAlYSi could be used.
  • the inserted material could advantageously be a layer of anti-wear material deposited on the upstream part 401 of the box 40 in the bearing zone at the other end 52 with the shock absorbing element 5 . Making a rough deposit in this manner changes the coefficient of friction and improves energy dissipation.
  • the shock absorbing element 5 is composed of at least two angular sectors fixed end to end and making up the complete annular shape of the shock absorber.
  • a minimum of two angular sectors satisfies assembly and differential expansion constraints encountered at the attachment area at the one end 51 with the annular support 24 of the high pressure turbine.
  • the number of angular sectors can be increased at will. For example, there may be two, six or eighteen angular sectors fixed end to end and making up the complete annular shape of the shock absorber. Eighteen is a particularly advantageous number of identical angular sectors, because it enables each of them to be attached to the annular support through screws/nuts 29 that arc also used to attach axial stop parts of the spacers. These parts are usually called stop plates.
  • the number of angular sectors and the number of blades must be a multiple of the number of attachment screws so that sectors can be identical.
  • any means of making the angular sectors by which they can be attached by existing screw/nut systems 29 to fix the axial stop parts is advantageous because the invention does not require any additional means of fixing the shock absorbing element.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US12/994,969 2008-05-28 2009-05-25 High pressure turbine of a turbomachine with improved assembly of the mobile blade radial clearance control box Active 2030-11-30 US8662828B2 (en)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
FR0853471A FR2931872B1 (fr) 2008-05-28 2008-05-28 Turbine haute pression d'une turbomachine avec montage ameliore du boitier de pilotage des jeux radiaux d'aubes mobiles.
FR0853471 2008-05-28
PCT/EP2009/056279 WO2009144191A1 (fr) 2008-05-28 2009-05-25 Turbine haute pression d'une turbomachine avec montage ameliore du boitier de pilotage des jeux radiaux d'aubes mobiles

Publications (2)

Publication Number Publication Date
US20110076135A1 US20110076135A1 (en) 2011-03-31
US8662828B2 true US8662828B2 (en) 2014-03-04

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US12/994,969 Active 2030-11-30 US8662828B2 (en) 2008-05-28 2009-05-25 High pressure turbine of a turbomachine with improved assembly of the mobile blade radial clearance control box

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US (1) US8662828B2 (fr)
EP (1) EP2300688B1 (fr)
JP (1) JP5615808B2 (fr)
CN (1) CN102046926B (fr)
BR (1) BRPI0912994B1 (fr)
CA (1) CA2725864C (fr)
FR (1) FR2931872B1 (fr)
RU (1) RU2503822C2 (fr)
WO (1) WO2009144191A1 (fr)

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20230146084A1 (en) * 2021-11-05 2023-05-11 General Electric Company Gas turbine engine with clearance control system
FR3140115A1 (fr) * 2022-09-22 2024-03-29 Safran Aircraft Engines Pièce d'amortissement des déformations d'un carter de récupération d'huile, ensemble qui le comporte et turbomachine ainsi équipée

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Publication number Priority date Publication date Assignee Title
FR2921410B1 (fr) * 2007-09-24 2010-03-12 Snecma Organe de verrouillage de secteurs d'anneau sur un carter de turbomachine, comprenant des moyens permettant sa prehension
JP2016507695A (ja) * 2013-02-08 2016-03-10 ゼネラル・エレクトリック・カンパニイ 吸引によるアクティブクリアランス制御システム
WO2014160953A1 (fr) * 2013-03-28 2014-10-02 United Technologies Corporation Joint étanche à l'air mobile pour turbine à gaz
FR3038655B1 (fr) 2015-07-06 2017-08-25 Snecma Ensemble comprenant un carter rainure et des moyens de refroidissement du carter, turbine comprenant ledit ensemble, et turbomachine comprenant ladite turbine
JP6563312B2 (ja) * 2015-11-05 2019-08-21 川崎重工業株式会社 ガスタービンエンジンの抽気構造
FR3081027B1 (fr) * 2018-05-09 2020-10-02 Safran Aircraft Engines Turbomachine comportant un circuit de prelevement d'air
KR102579798B1 (ko) * 2018-10-15 2023-09-15 한화에어로스페이스 주식회사 터보기기
FR3096395B1 (fr) * 2019-05-21 2021-04-23 Safran Aircraft Engines Turbine pour une turbomachine, telle qu’un turboréacteur ou un turbopropulseur d’avion

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JPH05214962A (ja) 1991-11-04 1993-08-24 General Electric Co <Ge> ガスタービンエンジンケーシング用熱制御装置及びガスタービンエンジン
US5772400A (en) 1996-02-13 1998-06-30 Rolls-Royce Plc Turbomachine
EP0892153A1 (fr) 1997-07-18 1999-01-20 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "Snecma" Dispositif d'échauffement ou de refroidissement d'un carter circulaire
JP2004003492A (ja) 2002-05-31 2004-01-08 Mitsubishi Heavy Ind Ltd ガスタービン圧縮機、及び、ガスタービン圧縮機のクリアランス制御方法
US6726446B2 (en) 2001-01-04 2004-04-27 Snecma Moteurs Stay sector of stator shroud of the high-pressure turbine of a gas turbine engine with clearance control
US20040251639A1 (en) * 2003-06-12 2004-12-16 Siemens Westinghouse Power Corporation Turbine spring clip seal
EP1505261A1 (fr) 2003-08-06 2005-02-09 Snecma Moteurs Dispositif de contrôle de jeu dans une turbine a gaz
US20070012043A1 (en) * 2005-07-18 2007-01-18 Siemens Westinghouse Power Corporation Turbine spring clip seal
US7287955B2 (en) 2004-01-16 2007-10-30 Snecma Moteurs Gas turbine clearance control devices
US7367776B2 (en) 2005-01-26 2008-05-06 General Electric Company Turbine engine stator including shape memory alloy and clearance control method
US20090104026A1 (en) 2007-10-22 2009-04-23 Snecma Control of clearance at blade tips in a high-pressure turbine of a turbine engine

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RU2133384C1 (ru) * 1995-11-22 1999-07-20 Открытое акционерное общество "Самарский научно-технический комплекс им.Н.Д.Кузнецова" Статор многоступенчатой турбомашины
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DE2258719A1 (de) 1971-12-09 1973-06-14 Westinghouse Electric Corp Gasturbine
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US5772400A (en) 1996-02-13 1998-06-30 Rolls-Royce Plc Turbomachine
EP0892153A1 (fr) 1997-07-18 1999-01-20 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "Snecma" Dispositif d'échauffement ou de refroidissement d'un carter circulaire
US6035929A (en) 1997-07-18 2000-03-14 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "Snecma" Apparatus for heating or cooling a circular housing
US6726446B2 (en) 2001-01-04 2004-04-27 Snecma Moteurs Stay sector of stator shroud of the high-pressure turbine of a gas turbine engine with clearance control
JP2004003492A (ja) 2002-05-31 2004-01-08 Mitsubishi Heavy Ind Ltd ガスタービン圧縮機、及び、ガスタービン圧縮機のクリアランス制御方法
US20040251639A1 (en) * 2003-06-12 2004-12-16 Siemens Westinghouse Power Corporation Turbine spring clip seal
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US7367776B2 (en) 2005-01-26 2008-05-06 General Electric Company Turbine engine stator including shape memory alloy and clearance control method
US20070012043A1 (en) * 2005-07-18 2007-01-18 Siemens Westinghouse Power Corporation Turbine spring clip seal
US20090104026A1 (en) 2007-10-22 2009-04-23 Snecma Control of clearance at blade tips in a high-pressure turbine of a turbine engine

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Search Report issued Jan. 16, 2009 in French Patent Application No. FR 709182 (with English translation of category of cited documents).

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20230146084A1 (en) * 2021-11-05 2023-05-11 General Electric Company Gas turbine engine with clearance control system
US11788425B2 (en) * 2021-11-05 2023-10-17 General Electric Company Gas turbine engine with clearance control system
FR3140115A1 (fr) * 2022-09-22 2024-03-29 Safran Aircraft Engines Pièce d'amortissement des déformations d'un carter de récupération d'huile, ensemble qui le comporte et turbomachine ainsi équipée

Also Published As

Publication number Publication date
CN102046926B (zh) 2014-11-05
JP5615808B2 (ja) 2014-10-29
CA2725864A1 (fr) 2009-12-03
US20110076135A1 (en) 2011-03-31
EP2300688B1 (fr) 2013-08-28
BRPI0912994B1 (pt) 2019-11-26
CN102046926A (zh) 2011-05-04
EP2300688A1 (fr) 2011-03-30
RU2010153596A (ru) 2012-07-10
WO2009144191A1 (fr) 2009-12-03
RU2503822C2 (ru) 2014-01-10
CA2725864C (fr) 2016-07-12
BRPI0912994A2 (pt) 2015-10-13
JP2011522150A (ja) 2011-07-28
FR2931872A1 (fr) 2009-12-04
FR2931872B1 (fr) 2010-08-20

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